1. Technical Field
The present invention generally relates to energy absorbing structural panels for aircraft and the like, and more particularly, to a variable density stitched-composite structural element.
2. Discussion
Modern military and civilian aircraft, and other types of vehicles, commonly include components which are designed to provide enhanced protection of the vehicle occupants during a crash. The ultimate goal of such crash protective components is to minimize the number of injuries to the occupants in survivable crash impacts. A secondary goal of such crash protective components is to maximize the structural integrity of the vehicle during normal or non-impact operation.
Aircraft design for crashworthiness involves a systems approach. In this approach, all of the aircraft components combine to yield crash protection for the occupants. For example, the fuselage structure, landing gear, and seats work together during a crash to absorb the aircraft kinetic energy. Ideally, the system operates to slow the occupants to rest without injurious loading during a crash. The energy-absorbing structure in an aircraft, generally located below the fuselage floor, is a very important part of this crash energy absorption system. In fact, this structure may absorb up to 8% of the aircraft kinetic energy during a crash.
While early generation crash energy absorption systems included mostly metal components, composite materials are now being applied to aircraft fuselage structures. Such composite materials offer potentially significant weight and cost reductions relative to metallic structures. However, composite materials are typically brittle and do not exhibit either plasticity or high elongation prior to failure. This may compromise the crashworthy design of an aircraft employing such composite components.
To enhance the energy absorption of composite components, a number of design solutions have been proposed. These design solutions include various crush initiators as well as ply drop-offs to sustain the crushing process after initiation. The drawback with these existing design solutions is that they are difficult to incorporate into the manufacturing process without significant cost increases.
Recently, stitched-composite materials have been applied to aircraft fuselage structures to offer additional benefits over non-stitched composites. Such benefits include superior damage tolerance and compression-after-impact capabilities. These improvements are provided due to the presence of aramid or glass thread stitches through the thickness of the laminates making up the composite structure.
Unfortunately, a drawback of stitched-composite structures is that the presence of the stitching may inhibit delamination which is part of the crushing process. Further, crush initiators and ply drop-offs which have been successfully incorporated into non-stitched composites may not work with stitched composites since the presence of the stitches significantly affects initiation and progression of the energy absorption process. Unless special design approaches are taken, stitched-composite subfloor structures may start to crush at higher loads that can subsequently precipitate unstable crushing due to elastic instability of the structure. This may result in reduced energy absorption.
In view of the foregoing, it would be desirable to provide a stitched-composite structural element for enabling a stable crushing process that starts at lower load levels, does not precipitate elastic instability, and ensures energy absorption by progressive delamination, bending, and fracture of the composite plies. This would result in load-crushing characteristics that maximize the energy absorption of the structural element.
The above and other objects are provided by an energy absorbing stitched-composite structural element. The element includes a composite substrate having a glass thread stitched therethrough. The stitching pattern of the thread varies in density from a low-density portion at an initial impact portion of the panel to a high density portion at a final impact portion of the panel. Preferably, the stitch pattern includes a plurality of rows laterally spaced apart relative to the height of the panel. The spacing between adjacent rows progressively decreases from the initial impact portion of the panel to the final impact portion of the panel. Alternatively, or in combination therewith, the number of stitches in each row varies from the fewest number of stitches proximate the initial impact portion of the panel to the greatest number of stitches proximate the final impact portion of the panel. In a highly preferred embodiment of the present invention, the stitched composite panel is incorporated into the subfloor structure of an aircraft fuselage.